The present invention generally relates to systems and methods for mounting an aircraft engine to an aircraft. More particularly, this invention relates to a mounting system and method adapted to reduce backbone deflection that can occur in an aircraft engine as a result of aerodynamic and thrust loads during aircraft operation.
FIG. 1 schematically represents a high-bypass turbofan engine 10 of a type known in the art. The engine 10 is schematically represented as including a nacelle 12 and a core engine (module) 14. A fan assembly 16 located in front of the core engine 14 includes a spinner nose 20 projecting forwardly from an array of fan blades 18. The core engine 14 is schematically represented as including a high-pressure compressor 22, a combustor 24, a high-pressure turbine 26 and a low-pressure turbine 28. A large portion of the air that enters the fan assembly 16 is bypassed to the rear of the engine 10 to generate additional engine thrust. The bypassed air passes through an annular-shaped bypass duct 30 between the nacelle 12 and an inner core cowl 36, and exits the duct 30 through a fan exit nozzle 32. The core cowl 36 defines the radially inward boundary of the bypass duct 30, and provides an aft core cowl transition surface to a primary exhaust nozzle 38 that extends aftward from the core engine 14. The nacelle 12 defines the radially outward boundary of the bypass duct 30, and the bypassed fan air flows between bypass duct flow surfaces defined by the nacelle 12 and core cowl 36 before being exhausted through the fan exit nozzle 32.
The nacelle 12 is typically composed of three primary elements that define the external boundaries of the nacelle 12: an inlet assembly 12A located upstream of the fan assembly 16, a fan cowl 12B interfacing with an engine fan case 42 that surrounds the fan blades 18, and a thrust reverser assembly 12C located aft of the fan cowl 12B. The thrust reverser assembly 12C comprises three primary components: a translating cowl 34A mounted to the nacelle 12, a cascade 34B schematically represented within the nacelle 12, and blocker doors 34C adapted to be pivotally deployed from stowed positions shown in FIG. 1 as radially inward from the cascade 34B. The fore end of each blocker door 34C is pivoted into engagement with the inner core cowl 36 when the door 34C is fully deployed, and as such the inner core cowl 36 of the core engine 14 is also part of the thrust reverser assembly 12C.
When installed on an aircraft, the engine 10 is supported by an aircraft structure, for example, a pylon (not shown) that extends outward from the aircraft. In the case of an engine mounted to a wing, the pylon typically extends downwardly beneath the wing. Structural components of the pylon are connected to a frame of the core engine 12 that supports the rotating components of the compressor 22 and turbines 26 and 28. The engine frame typically includes a forward frame adjacent the compressor 22, an aft frame adjacent the turbines 26 and 28, and an engine casing that connects the forward and aft frames. The engine casing is often referred to as the backbone of the engine 10. Aircraft engines of the type represented in FIG. 1 are typically mounted and secured to an aircraft in two planes normal to the engine centerline 40. One mount is typically connected to the forward frame often just rearward of the fan assembly 16, and a second mount is typically connected to the aft frame near the turbine section.
During climb and certain aircraft maneuvers, the centerline 40 of the engine 10 is pitched relative to the direction of approaching airflow, with the result that the nacelle 12 can be subjected to upward aerodynamic loading. This aerodynamically-induced load, often referred to as the inlet load and represented by the vector Fi in FIG. 1, is in addition to the thrust load, represented by the vector Ft in FIG. 1. These loads induce bending moments in the engine casing (backbone), with the result that the backbone is deflected (bends) from its concentric position about the engine centerline 40. Maintaining concentricity of the engine backbone about the centerline 40 is important from the standpoint of minimizing blade tip clearances within the compressor 22 and turbine sections 26 and 28 of the engine 10, which has the beneficial effect of improving engine specific fuel consumption (SFC) and fuel burn. In addition, reduced backbone bending reduces the incidence of blade tip rub encounters with the surrounding engine structures (including the fan case 42), which promotes in-service performance retention. Engines with a longer interval for time on-wing to removal for service provide reduced service contract costs to their operators.
Approaches for reducing backbone deflection in high-bypass turbofan engines have included reinforcement of the engine frame. However, such approaches typically increase weight and cost and may not be entirely effective, particularly as inlet and thrust loads increase with larger nacelles and higher thrusts. Other approaches have included orienting the forward mount plane to move its focal point toward the engine centerline and forward toward the inlet load vector (Fi). However, this approach has not entirely eliminated backbone bending especially in flight regimes where maximum inlet loads are encountered, such as when an aircraft rotates during take-off. Consequently, there is an ongoing need for approaches capable of reducing backbone deflection in high-bypass turbofan engines.